Terminal guidance method and a guided missile operating according to this method

ABSTRACT

A guidance method is provided for the terminal portion of the trajectory of a guided missile having a sensor and comprising two sections coupled together by a central shaft and free to rotate with respect to one another about the longitudinal axis of the missile; one section comprises a drive means for controlling the roll attitude of this section and a gas generator which feeds a nozzle for providing a transverse throat force and the other section has a stabilizing tail unit formed by a set of fins able to be opened out.

BACKGROUND OF THE INVENTION

The invention relates to guided missiles and, more precisely, to amethod for guiding a missile, applicable during the terminal portion ofthe flight path; it also relates to a guided missile operating accordingto this guidance method.

There exists a demand for AIR to GROUND missiles capable of stopping, atrelatively large distances, the threat presented by land formationsformed more especially by motorized vehicles such as armored vehiclesadvancing in groups over the terrain. These armored vehicles, by theirnature, radiate thermal energy and thus constitute potential targetswhich may be detected and located by a missile fitted for example withan electro-optical E.O. sensor operating in the IR band of theelectromagnetic spectrum. Furthermore, the missile may be provided witha military charge capable of piercing the protecting armor of armoredvehicles. It is possible to direct the firing of such a missile towardsa group of armored vehicles; however, there remains the problem ofsupplying, during the terminal portion of the downward trajectorytowards the ground, the trajectory corrections for providing impact of amissile on one of the vehicles detected by the EO sensor.

A missile is already known comprising guidance means for correcting, inthe terminal phase of the flight path, the possible error between thedirection of a target and the direction of impact of the missile on theground, in free fall. To this end, the base of this missile of the priorart is equipped with a set of fins which impart to the body of themissile a self rotating movement at a substantially constant angularspeed about its longitudinal axis. In the head of the missile isdisposed an electro-optical EO sensor and, finally, in the middle partof the body a lateral impeller may supply a predetermined thrust forcewhose direction is normal to the speed vector of the missile. The EOsensor is formed by a plurality of photodetector cells arranged in aring in a plane perpendicular to the axis of the missile, so as toprovide a hollow conical field of view. Thus, the surface of the groundcovered by the field of view of the EO sensor is gradually reduced as afunction of the decreasing altitude of the trajectory. When the targetcomes into the field of view of the sensor, its image falls on one ofthe photodetector cells which determines, in polar coordinates, theposition of the target with respect to the orientation of the impeller.The output signal of the EO sensor is used to supply an order fortriggering the lateral impeller at the moment when the orientation ofthis latter is opposite the direction of the detected target.

This missile of a relatively simple prior art construction does notallow the degree of efficiency sought to be attained and, moreespecially, a probable hit on the target to be obtained. To attain thisaim, the guidance method proposed uses a sensor for tracking the targetwhich measures the rotation of the missile-target line of sight.

SUMMARY OF THE INVENTION

According to one aspect of the invention there is provided a method forguiding a missile during a terminal portion of the missile's trajectory.The missile has a sensor with a beam sensitive to energy radiated by apotential target. The method includes steps for seeking of a target, andsteps for piloting the missile. The target seeking steps include: (a)immobilizing the beam of the sensor along the longitudinal axis of themissile; (b) imparting to the missile a rotation about the longitudinalaxis of the missile at a given angular roll speed, and aspiral-line-movement about the trajectory of the missile, so that thebeam of the sensor describes a surface of revolution (by creating atransverse thrust force normal to the direction of the speed of movementof the missile), and (c) detecting an image of a possible target pickedup by the beam of the sensor. The steps for piloting the missile include(a) freeing the beam of the sensor and maintaining the axis of this beampointed at the image of the detected target to measure the rotation ofthe missile-target line of sight; (b) elaborating a piloting orderproportional to the measured magnitude of the rotation of the line ofsight; and (c) applying this piloting order to modify the roll attitudeof the missile.

According to a further aspect of the invention there is provided aguided missile with a sensor sensitive to energy radiated by a potentialtarget. The missile has first and second main sections which aremutually coupled together and which rotates with respect to each otherabout the longitudinal axis of the body of the missile. The firstsection has a sensor, and includes a drive with a first member integralwith the structure of the front (first main) section. A second member isphysically coupled to the second main section. A control input isconnected through an amplifier to a generator-of-piloting-orders so asto vary the roll attitude of the body of the missile. A gas generatorfeeds a lateral nozzle to provide a transverse thrust force. The secondmain section includes at its base a stabilizing tail unit formed of finswhich may be folded out. The sensor of the first section is providedwith a locking device for fixing a beam along the longitudinal axis ofthe missile while the missile is seeking a target.

Another object of the invention consists in conferring on the missile agiven initial speed of movement along its trajectory and maintaining itsubstantially constant along the trajectory.

Another object of the invention is to vary the angular speed ofself-rotation of the body of the missile along its terminal trajectory.Furthermore, the second member of the drive means is coupled to the rearsection of the missile by a central shaft.

According to a further object of the invention the rear section of themissile comprises a compartment for housing a releasable brakingparachute for reducing the ballistic speed of the missile over theportion of the trajectory preceding the terminal phase.

BRIEF DESCRIPTION OF THE DRAWINGS

The characteristics and advantages of the invention will be clear fromthe detailed description which follows, made with reference to theaccompanying drawings which illustrate the guidance method and oneembodiment of the guided missile; in these drawings:

FIG. 1 shows a guided missile of the prior art,

FIG. 2 shows the method of constructing the electro-optical sensor ofthe missile of the prior art,

FIG. 3, in a simplified schematical form, shows a guided missilecomprising the means required by the guidance method of the invention,

FIG. 4 shows a cross sectional view of the guided missile of FIG. 3,

FIG. 5 is a plane diagram of axes x,z associated with the ground andindicating the principal parameters which determine the extent of theground swept by the beam of the sensor,

FIG. 6 is a diagram of a trihedron x, y, z associated with the groundand illustrating the method of searching for a potential target,

FIG. 7 shows a detailed view of a portion of the trajectory of themissile,

FIG. 8 is a simplified diagram showing a variation of the seekingtrajectory,

FIG. 9 shows the law of acceleration conferred on the missile as afunction of the magnitude of the rotation of the missile-target line ofsight,

FIG. 10 illustrates the law for controlling the roll attitude of thebody of the missile as a function of the magnitude of the rotation ofthe missile-target line of sight.

FIG. 11 is a longitudinal section of a guided missile according to theinvention,

FIG. 12 shows, in an exploded view, the elements of an electric torquermotor,

FIG. 13 shows one embodiment of the stabilizing tail unit,

FIG. 14 illustrates one application of the guided missile to thedestruction of a group of land vehicles,

FIG. 15 is an exploded view of the compartment of a carrier projectilehousing a plurality of missiles,

FIG. 16 is a sectional view of the carrier projectile showing therelative arrangement of the guided missiles in the compartment,

FIG. 17 is a diagram of the components of the rotational vector of themissile-target line of sight in an absolute trihedron and in the missiletrihedron,

FIG. 18 shows, in the form of a block diagram, the elements of the servoloop for tracking the missile.

FIG. 1 shows, in a simplified form, the missile of the prior artmentioned in the preamble of this application as well as thecorresponding terminal guidance method. Missile 1 is equipped with a setof fins 2 whose configuration imparts to the body of this projectile anangular speed of self-rotation ω_(r) about its longitudinal axis Xcarrying the speed vector V of movement of the projectile along itstrajectory. In free fall, the trajectory of the missile is inclined byan angle θ_(t) and this missile strikes the ground at a point 4 offsetangularly by an angle θ_(c) from a potential target 6.

For modifying the trajectory of the missile, the missile is fitted witha lateral impeller 3 and an electro-optical sensor 5 which supplies asignal for triggering this impeller, this triggering signal resultingfrom the measurement of the error angle θ_(c). The result is that thespeed vector V of the projectile is modified by an amount V_(c) toprovide a resulting speed vector V_(r) offset by the angle θ_(c) fromthe speed vector V to obtain impact of the missile on the target.

FIG. 2 shows the embodiment of the electro-optical sensor 5 carried bythe missile 1 described in FIG. 1. This EO sensor is formed essentiallyby a plurality of photo-conducting elements 7 arranged in a ring in aplane orthogonal to the longitudinal axis X of the body of the missileto supply a predetermined hollow conical field of view with angularaperture θ and angular width β (FIG. 1). When the image 8 of target 6 isdetected by one of the photoconducting elements 7, such as element7_(i), the width of the relative angle A between the direction of theimpeller 3 and the photoconducting element 7_(i) is measured by the EOsensor and fed to a computing circuit which determines the moment fortriggering the impeller 3 corresponding to the impeller passing into thedirection of the detected target.

FIG. 3 shows, in a simplified schematical form, a guided missile 3 whichcomprises means specific to the terminal guidance method of theinvention. This missile comprises: a sensor 11, sensitive to the energyradiated by a potential target, situated in the head of the missile, ameans 12 for providing a transverse thrust P_(o) passing through thecenter of gravity G of the missile and a means 13 for controlling theroll attitude of the body of missile 10 about its longitudinal axis X.The sensor is provided with a locking means for immobilizing its beamalong the longitudinal axis X, means for detecting the possible presenceof a target intercepted by this beam and angular tracking means formeasuring the rotation η of the target-missile line of sight (L.O.S.).The means 12 for providing a transverse thrust P_(o) comprises acombustion chamber which supplies a lateral nozzle whose thrustdirection is inclined, by an angle α, to the longitudinal axis X of themissile; the result is that the transverse F_(N) and longitudinal F_(L)components of the force F applied to the missile are given by thefollowing relationships:

    F.sub.N =F cos α

and

F_(L) =F sin α

to which correspond the normal acceleration γ_(N) given by the followingrelationship ##EQU1## and the longitudinal acceleration γ_(L) given bythe following relationship: ##EQU2## where M is the mass of the missileand g the magnitude of the Earth's field of gravity.

FIG. 4 shows a section of the missile 10, with axes X, Y and Z; andshows the components F_(Y) and F_(Z) of the normal force F_(N) as afunction of the roll angle φ of the body of the missile about itslongitudinal axis X. These components F_(Y) and F_(Z) are given by thefollowing relationships:

    F.sub.Y =F.sub.N cos φ

    F.sub.Z =F.sub.N sin φ

The body of the missile may rotate in both directions, with respect toaxis X at an instantaneous angular speed φ. The magnitudes φ and φ maybe measured on board the missile and used respectively for controllingthe roll attitude and the self-rotational speed of the body of themissile.

FIG. 5 is a plane diagram with axis x, z associated with the ground inwhich are shown the principal parameters which determine the extent ofthe ground swept by the beam 14 of the EO sensor carried by thepreviously described missile 10. The center of gravity G of the missileis driven at a speed of movement V directed along the longitudinal axisX of the body of the missile and it is subjected to a system of threeforces. A first force, normal corresponds to an acceleration γ_(N)normal to the speed vector V, a second force, longitudinal correspondsto an acceleration directed along the longitudinal axis X, and a thirdforce, of the Earth's gravity to which corresponds the accelerationvector g directed along the vertical of the locality. The beam 14 of themissile has a relatively narrow half aperture angular field ε, a fewdegrees for example. The straight line G.I. of the downward trajectoryof the missile is inclined by an angle θ_(o) with respect to thehorizontal. Since the body of a missile is subjected to aself-rotational speed φ about its longitudinal axis X and since the beam14 of the EO sensor is immobilized along this longitudinal axis X, theresult is that the beam 14 describes as a function of time a hollowcone, which is a surface of revolution, with axis GI whose external andinternal half apertures have for respective values (θ+ε) and (θ-ε).Since the altitude R_(h) of the missile above the ground is reducedproportionally to the time, the axis 15 of beam 14 describes on theground, as a function of time, a converging spiral with radius R_(s)centered on point I. The extent of the surface of the ground swept bybeam 14 is a circle when the descent angle is equal to 90° and anellipse of small eccentricity when the value of this angle θ remainshigh, 60° to 70° for example.

FIG. 6 is a diagram in a trihedron x, y, z, associated with the groundwhich illustrates the method for seeking a target by means of themissile described previously, in a particular case corresponding to adescent angle θ_(o) equal to 90°. We will consider here the case wherethe rotational speed φ of the missile about its longitudinal axis X ismaintained constant as well as the speed V of the missile while ignoringthe force of resistance of the air and considering that the longitudinalacceleration force γ_(L) produced by the nozzle of the missile and theforce of gravity g are equal and opposite values. The trajectory S fromthe center of gravity G of the missile describes a helix carried by acylinder 16 with vertical axis z passing substantially through point Iand the radius of this cylinder has a magnitude r. The extent A_(s) ofthe surface of the ground swept by the beam 14 of the EO sensor,describing a surface of revolution, is given by the following formula:

    ΔA.sub.s =π·(R.sub.h ·tan (θ+ε)).sup.2

The surface of the ground ΔA_(s) intercepted by the optical beam is anellipsis in which the magnitudes of the axes ΔR_(s) and ΔR'_(s) aregiven respectively by the relationships: ##EQU3##

The oblique distance R_(d), between the missile and the surface ΔA_(s)of the ground intercepted by the beam of the EO sensor is given by thefollowing relationship: ##EQU4##

The horizontal distance R_(s) between the point I and the surface ΔA_(s)is given by the following relationship:

    R.sub.s =R.sub.h ·tan θ

In FIG. 6, there is also shown a target c driven at a speed V_(c) anddistant from point I by a value R_(c). To ensure a high probability ofdetecting a target such as c, the angular speed Ω of the beam 14 of theEO sensor must be determined so as to obtain a certain amount ofoverlapping of the successive sweep frames.

The passing time of the optical beam over a target C is given by thefollowing relationship: ##EQU5## where Ω is the angular rotational speedof the beam about the vertical axis z.

FIG. 7 shows a detailed view of a portion of the trajectory S of missileshown in the preceding figure. The speed vector V of the missileoriginates at point G representing the center of gravity of the missile,this speed vector V is contained in a plane P tangent to a generatrix ofa cylinder 16 carrying point G. The components of the speed vector V arethe vertical component V_(h) and the orthogonal component V_(t) given bythe following relationships:

    V.sub.h =V·cos θ

and

    V.sub.t =V·sin θ

The speed component V_(t) is tangent to the circle having a center O anda radius r. From the general relationships of the dynamics ##EQU6## withΩ=φ/cos φ. By combining the preceding relationships, we obtain the valueof the inclination angle θ of the speed vector V of the missile withrespect to the generatrix C.I. of the cylinder ##EQU7##

FIG. 8 is a simplified diagram showing a variation of the method forseeking a target on the ground. According to this variation, the angularroll speed φ of the missile about its longitudinal axis X, is varied asa function of the the altitude R_(h) of the missile above the ground.The preceding formulae giving the values of the width ΔR_(s) of thesuccessive sweep frames and the angle of inclination θ of the speedvector V of the missile may be rewritten in an approximate form:

    ΔR.sub.s =2H·ε meters ##EQU8## assuming that the values of angles ε and θ are still small.

It follows, that if the adjacent sweep frames of the beam of the EOsensor overlap with an overlapping factor of 50%, we have the followingrelationship: ##EQU9##

The result is that the trajectory S of the center of gravity G of themissile is inscribed on the surface of a cone of radius r such that:##EQU10##

We have just analysed in detail the initial portion of the terminaltrajectory of the missile corresponding to the phase of seeking apossible target situated in a zone A_(s) on the ground centered on thedescent axis of the missile. In what follows, the final portion of thetrajectory of the missile will be described corresponding to theacquisition of the image of the target by the sensor and, consecutively,to piloting the missile so as to obtain impact on the detected target.Referring again to FIGS. 6 and 7, it can be seen that, when plane P, inits rotational motion with respect to the vertical axis z passes, at agiven moment, in the vicinity of point C corresponding to the positionof a target and that the following relationship:

    R.sub.c ≃R.sub.h ·tan θ

is substantially satisfied, the EO sensor detects the image of thetarget. From this moment, the EO sensor supplies the following outputsignals: a first output signal indicating the presence of a target inbeam 14 and a second output signal proportional to the rotational speedof the missile-target line of sight. The first output signal is used forfreeing the beam of the optical sensor and allowing angular tracking ofthe sensor on the image of the target; the second output signal, oncethe angular tracking has been ensured, is fed to a computing means forcontrolling the roll attitude of the body of the missile and,consequently, directionally piloting the missile.

FIG. 9 is a diagram which shows the rotational speed vector η of themissile-target line of sight. F_(N) being the thrust force normal to thespeed vector V passing through the longitudinal axis X of the missileand Δφ the orientation angle of this thrust force F_(N).

The equation of the piloting law of the missile is in the form

    γ.sub.η =γ.sub.N cos Δφ=2.sub.η ·V+A(η-η.sub.o)·V

which corresponds to a law of proportional navigation. The gain Acomprises a bias η. If, by way of example, we make the acceleration(γN)/2 correspond to this bias, which has the advantage of giving anequal margin of maneuverability on each side of the magnitude η_(o)given by the following relationship: ##EQU11##

Consequently, the input piloting signal is proportional to the magnitudeη and the response is the magnitude Δφ of the orientation of the thrustforce F_(N) with respect to the direction of the rotational vector ηsuch that

    Δφ=Arc cos (Kη+K.sub.o)

since the terms η_(o) and V of the equation of the guidance law areconstants.

FIGS. 9 and 10 shown facing each other illustrate the laws of theacceleration γ and of the roll piloting angle Δφ of the missile as afunction of the modulus of the rotational vector η.

FIG. 17 is a diagram showing the components of the rotational vector ηin an absolute trihedron U,V and in the missile trihedron Y,Z referencedto the direction of the piloting nozzle.

FIG. 18 shows, in the form of a block diagram, the servo loop fortracking the missile, which comprises the following elements: theguidance sensor 100 which delivers the components η_(y) and η_(z) of therotational vector of the missile-target line of sight, these twocomponents are fed to a resolver device 110 and an operator 120 whichelaborates the modulus of the rotational vector |η|, this modulus |η| isapplied to an operator 130 for supplying an output signal Δφ inaccordance with the guidance law shown in FIG. 10 and, through a servomotor 140, rotates the resolver 110 through an equivalent angle;finally, the output signal V.sub.ε is applied to the roll control means150 of the missile body.

The crossed component of the acceleration γ_(T) =γ_(N) sin Δφ generatesa spiral movement of the interception trajectory of the missile. Theangular roll speed φ of the body of the missile is then given by thefollowing relationship: ##EQU12## in which V_(R) is the relative speedand R_(d) the remaining missile-target distance. The result is that theacceleration component γη ensures biased proportional navigation and theacceleration component γ_(T) generates a spiral trajectory but has noeffect on the convergence of the guidance on to the target.

The guidance method which has just been described may be applied to aguided missile of moderate caliber, for example of the order of 100 mm,and the magnitudes of the main parameters enumerated above may, by wayof indication, be situated about the following values: speed of movementV of the missile along its trajectory of the order of 50 ms⁻¹, angle ofdescent θ_(o) between 60° and 90°, angle of inclination θ of the missilespeed vector with respect to the descent axis between 10° and 15°,angular half aperture ε of the beam of the sensor of the order of 4° to8°, altitude R_(h) of the missile at the time of igniting the gasgenerator, of the order of 500 m. For these values of the mainparameters, the travel duration of the terminal portion of thetrajectory is between 10 and 15 seconds and, for a normal accelerationvalue γ_(N) of the order of 25 ms⁻², the angular rotational speed duringrolling φ is of the order of 2.5 rad.s⁻¹, the surface of the groundswept by the beam of the sensor is about 5.10⁴ m². All the values ofthese parameters may vary depending on the specific mission of themissile.

FIG. 11 is a view along a longitudinal section of one embodiment of aguided missile operating in accordance with the guidance method whichhas just been described.

The guided missile 10 comprises two main sections; a first main section20, called "front section", and a second main section 30, called "rearsection", which rotate with respect to each other about the longitudinalaxis X of the missile. The front and rear sections are mutually coupledtogether through a central shaft 21. This shaft is rigidity locked withthe rear section, and is carried by two bearings 22a and 22b inside thefront section. Inside the front section 20 are disposed the followingelements:

an EO sensor 23 situated behind a transparent dome 23a,

a drive means 24 for controlling the roll attitude of this frontsection; this drive means comprising: a first member 24 integral withthe mechanical structure of this front section and a second member 24bphysically coupled to the central shaft 21 coupling together the frontand rear sections of the missile,

a compartment 25 containing the electronic circuits associated with theEO sensor on the one hand and with the drive means 24 on the other and,

a gas generator 26 coupled to a lateral nozzle 27 whose output orificeis situated on the external lateral wall of this front section.

The rear section 30 of the missile, physically integral with the centralcoupling shaft 21 is provided, at its base, with a stabilizing tail unit31 formed by a set of unfoldable fins 32; in this figure, only two finshave been shown; one of the fins 32a is shown in the unfolded or activeposition whereas the other fin 32b is shown in the folded or inactiveposition. Inside this rear section are disposed the following elements:

the military charge 33 of the missile and

a compartment 34 for storing a parachute 35 released on the trajectoryof the missile, then dropped during flight.

Such a missile may be characterized by its following principaldimensional parameters: its caliber equal to its external diameterD_(o), its overall length L_(o), the span of its fins L_(E) and itstotal mass M_(o).

The principal elements mentioned above will now be described. The EOsensor 23 is sensitive, for example, to the thermal energy radiated bythe vehicles to be intercepted and the dome 23a is transparent to thecorresponding IR radiation. This EO sensor comprises an optical assemblyat the focal point of which is disposed a photodetecting element 23c forproviding a beam 14 with half aperture equal to an amount ε, this beambeing materialized by its axis 15. The whole formed by the opticalassembly and the photodetecting element 23c is carried by a gyroscopecomprising locking means (tulipage) for immobilizing the axis of theoptical beam 14 along the longitudinal axis X of the missile andprecessional means for orientating, in the no locked position, thisoptical beam in space. Furthermore, this EO sensor comprises electronicmeans for detecting the presence of a thermal source intercepted by thebeam and means for latching the axis of the optical beam to the straightline between target and missile.

The drive means 24 for controlling the roll attitude of the frontsection of the missile is a torquer motor. A torquer motor is a rotarymultipolar electrical machine which may be coupled in direct drive withthe load to be driven. This type of machine transforms electric controlsignals into a sufficiently high mechanical torque to obtain a givendegree of precision in a speed or position servo system. A torquer motorof the "pancake type", because of its design, may be easily integratedin the structure of the missile. As shown in FIG. 12, this type oftorquer motor comprises essentially three elements: a stator 24a whichprovides a permanent magnetic field, a laminated wound rotor 24bintegral with a segmented collector 24c and a brush carrying ring 24dequipped with connections for receiving the control signals. Because ofits mechanical characteristics, this torquer motor ensures rigidcoupling with the load, resulting in a high mechanical resonancefrequency; because of its electrical characteristics, the intrinsicresponse time of a torquer motor may be short and its resolution high.Moreover, the torque delivered increases proportionally with the inputcurrent and is independent of the speed or of the angular position.Since the torque is linear as a function of the input current, this typeof machine is free of operating threshold. Torquer motors arecommercialized more particularly by the firms ARTUS (France) and INLAND(U.S.A.). The second member 24b of the drive means, because of itsconnection with the rear tail unit part of the missile, is subject to aresistant torque resulting from the combination of the inertial torqueof this rear section and from the aerodynamic torque provided by thetail unit. The first member 24a of the drive means comprises a controlinput which is connected to an amplifier which includes correctorelectric networks. The input of this amplifier, during the phase ofseeking a target by the sensor, receives an electric signal resultingfrom the comparison of the angular roll speed φ of the body of themissile and a reference value. The angular roll speed of the body of themissile may be provided by a rate gyro whose sensitive axis is alignedalong the longitudinal axis of the missile. The reference value may bevaried as a function of time, i.e. depending on the altitude of themissile above the ground. During the phase for piloting the missile onto the detected target, the input of the amplifier of the drive meansreceives an electric signal for controlling the roll attitude of thebody of the missile so as to cancel out the rotation of themissile-target line of sight.

The tail unit 31 of the missile is formed by fins movable between aposition folded back against the body of the missile and an activeunfolded or folded out position. Considering the relatively low movingspeed V of the missile, the tail unit is required to provide a highaerodynamic stabilizing torque, this is obtained by means of fins ofgreat extension which are laid tangentially against the body of themissile. FIG. 13 is a perspective view of the tail unit assembly, thefins situated at the front of the figure being omitted for the sake ofclarity. The body 31a of the tail unit is an annular part having, forexample, an inner thread 31b for fixing same to the base of the rearsection 30 of the missile. This annular part comprises a set of slopingfork-joints 31c spaced apart evenly around the periphery of the part. Inthese fork-joints, a slit 33 with parallel faces receives the hinginglug 34 of the fin 32 which, by by means of a pin, may pivot in holes 33aand 33b. From the mechanical point of view, the tail unit is completed,for each of the fins, by a device for locking it in the folded outposition. This device is formed, for example, by a spring lockingmechanism 36 which actuates a pin 37 which may engage in a lateral notchprovided for this purpose in the hinging lug of the fin. A detailedembodiment of this type of tail unit has been described in French patentPV No. 53 419, filed on Mar. 15, 1966 and published under the No. 1 485580. Besides its stabilizing function, the tail unit supplies anaerodynamic resistant torque which is transmitted to the second member24b of the drive means 24.

The gas generator 26 is essentially formed by a combustion chamberinside which are disposed two blocks 26a and 26b of solid propergol.Between these two blocks of propergol is located an ejection nozzle 27whose output orifice opens into the lateral wall of the body of themissile. The thrust direction of the gases Po is inclined by an angle αtowards the front of the missile so as to provide the two accelerationforce components: the longitudinal force F_(L) for compensating theforce of the Earth's gravity and the normal force F_(N) used incombination with the roll attitude of the body of the missile to varythe orientation of the speed vector V of the missile. The section of thecombustion chamber and, consequently, the section of the propergolblocks may be of a toric shape so as to leave a free passageway aboutthe longitudinal axis X of the missile, and more especially fordisposing the coupling shaft 21 of the front and rear sections of themissile.

The total mass m_(p) of propergol must satisfy the followingrelationship: ##EQU13## where F is the required thrust force, Td themaximum travel duration of the missile over the terminal portion of itstrajectory and Is the specific impulse of the propergol used.

The military charge may be advantageously of the so-called "hollowcharge" type which produces a jet capable of piercing the protectingarmor of vehicles. So as to ensure free passage of the jet along thelongitudinal axis of the missile, the shaft 21 for coupling the frontand rear sections of the missile together comprises a recess 21a in itsaxial portion; moreover, a free passage may be provided also in thecentral part of compartment 25 containing the electronic circuitsassociated with the EO sensor 23 and with the drive means 24.

The braking parachute 35 of the missile may be a parachute similar tothose used in the technique of braked projectiles such as aviationbombs. With this parachute are associated release and dropping devicesnot shown. The duration of the action of the parachute depends on themass Mo of the missile and on the ratio of the cruising speed to thepredetermined speed V over the terminal portion of the trajectory of themissile.

The guided missile which has just been described in detail may be amissile of average caliber of the order of 100 mm and with an elongationfactor of about 6 to 7 for a weight of 10 to 15 kgs. However, it may bepointed out that all its values may be modified within wide limitsdepending more particularly on the destructive power of the militarycharge carried.

The guided missile, in itself, such as has just been described, may forma sub-projectile of a larger sized projectile whose main function is tocarry this or a group of such sub-projectiles over the cruising portionas far as the terminal position of the firing trajectory.

Referring now to FIG. 14 which illustrates the transitory portionbetween the cruising portion and the terminal portion of the firingtrajectory, the carrier projectile 50 transports sub-projectiles orguided missiles 51, 52 and 53 situated in a section 54. On reaching thetransition portion of the trajectory, the guided missiles are ejectedand dispersed at a high initial speed substantially equal to that of thecarrier projectile and are at a predetermined altitude above the ground.So as to reduce their initial moving speed to reach the adequate speed Vfor the acquisition and interception of targets, the braking parachute35 of the missile is released for a determined period of time, afterwhich the mechanical connection between the missile and the parachute isbroken so as to drop this latter. The stabilizing tail unit 31 isunfolded and the front section of the missile is set in self-rotation.Then, the gas generator for producing the transverse thrust force F_(N)is activated and the phase for seeking a potential target situated onthe ground may begin. Because of the ejection force imparted by thecarrier vehicle 50 at the time of separation thereof from thesub-projectiles 51 to 52, there results a certain dispersion distanceR_(D) at the moment when the operation for seeking targets by the sensorof the sub-projectile begins.

FIG. 15 is a partial exploded view of section 54 of the carrierprojectile 50 which shows one example of installing a group of threeguided missiles 51, 52 and 53. These missiles are evenly spaced apartabout the longitudinal axis of the carrier projectile, and furthermore,an identical group of missiles may be installed in tandem, if necessary.

FIG. 16 is a cross section of the carrier projectile 50 which shows therelative arrangement of the guided missiles 51, 52 and 53 inside thehousing section 54. The guided missiles abut against elements 55actuated by an ejection mechanism 56 whose complementary function is tocommunicate a certain amount of movement to the missiles during ejectionthereof so as to ensure a predetermined relative dispersion. Theejection mechanism 56 may be of a known mechanical type actuated byhydraulic, pneumatic or possibly electric means. So as to minimize thecross section of the carrier projectile, the missiles may be providedwith a tail unit formed of four fins 32, capable of being folded out, soas to allow a certain material recessing thereof.

Table 1 is a recapitulary table of the sequence of the principaloperations effected by the missile during its firing trajectory.

The guided missile of the invention is not limited in itscharacteristics and applications to the embodiment described. Moreespecially, the sensor may be of the passive or semi-active type andoperate in the optical or radar bands of the electromagnetic spectrum,the relative arrangement of the elements such as the drive means 24 andthe military charge 23 may be modified.

The invention is not limited to its application to an independentmissile, but also applies to a missile carried by conventional vehiclesor aircraft.

                  TABLE 1                                                         ______________________________________                                        t.sub.o                                                                              end of the carried cruising phase of the missile,                             locking of the sensor to the longitudinal axis of the                         missile,                                                                      starting up of the rotor of the gyroscopio elements                           of the missile,                                                               setting of the gyroscopic references,                                         energization of the primary electric energy source,                    t.sub.o + T.sub.1                                                                    ejection of the missile from its carrier,                              t.sub.o + T.sub.2                                                                    opening of the braking parachute,                                      t.sub.o + T.sub.3                                                                    dropping of the braking parachute and opening of the                          stabilizing tail unit,                                                 t.sub.o + T.sub.4                                                                    ignition of the gas generator and application of a                            transverse thrust force to the missile and sensitiz-                          ation of the sensor of the missile,                                    t.sub.o + T.sub.5                                                                    the body of the missile set in self-rotation about                            its longitudinal axis,                                                 t.sub.o + T.sub.6                                                                    detection of the presence of a potential target on                            the ground and unlocking of the sensor and locking of                         the beam of the sensor on the image of the detected                           target,                                                                t.sub.o  + T.sub.7                                                                   measurement of the rotation of the missile-target line                        line of sight and elaboration of the order for pilot-                         ing the missile,                                                       t.sub.o + T.sub.8                                                                    impact on the target and setting off of the military                          charge.                                                                ______________________________________                                    

I claim:
 1. A method for guiding a missile during a terminal portion ofthe missile's trajectory, said missile having a sensor with a beam saidsensor being sensitive to energy radiated by a potential targetcomprising the following steps for seeking the target:(a) immobilizingthe beam of the sensor along the longitudinal axis of the missile; (b)imparting to the missile:a rotation about the longitudinal axis of themissile at a given angular roll speed, and a spiral line movement inorder that the beam of the sensor describes a surface of revolution bycreating a transverse thrust force normal to the direction of the speedof movement of the missile; (c) detecting an image of a possible targetpicked up by the beam of the sensor; and comprising the following stepsfor piloting the missile: (d) freeing the beam of the sensor andmaintaining the axis of this beam pointed at the image of the detectedtarget to measure the rotation of the missile-target line of sight; (e)elaborating a piloting order proportional to the measured magnitude ofthe rotation of the line of sight; and (f) applying this piloting orderto modify the roll attitude of the missile.
 2. The guidance method asclaimed in claim 1, wherein the speed of movement of the missile isestablished at a given value, at the moment when said missile enters theterminal portion of its trajectory.
 3. The guidance method as claimed inclaim 2, wherein the speed of movement of the missile over the terminalportion of the trajectory is maintained substantially constant bycreating a longitudinal thrust force having a magnitude substantiallyequal to the force resulting from the Earth's gravity field and in adirection aligned with the longitudinal axis of the missile.
 4. Theguidance method as claimed in claim 3, wherein the angular roll speed ofthe body of the missile is increased along the terminal portion of thetrajectory of the missile.
 5. A guided missile having a sensor which issensitive to energy radiated by a potential target and comprising firstand second main sections mutually coupled together and rotating withrespect to each other about the longitudinal axis of the body of themissile;the first section containing an electro-optical sensor andcomprising a drive means having a first member integral with thestructure of the first section, a second member physically coupled tothe second main section, and a control input connected through anamplifier to a generator of piloting orders so as to vary the rollattitude of the body of the missile, and a gas generator which feeds alateral nozzle so as to provide a transverse thrust force; the secondmain section comprising at its base a stabilizing tail unit formed offins able to be folded out, said sensor of the first section havinglocking device means for immobilizing a sensor beam along thelongitudinal axis of the missile during a step of seeking a target. 6.The missile as claimed in claim 5, wherein the second member of saiddrive means is mechanically coupled to the rear section of the missileby a central coupling shaft.
 7. The missile as claimed in claim 6,wherein said drive means is an electric torquer motor.
 8. The missile asclaimed in claim 7, wherein said rear section of the missile comprises amilitary charge of the "hollow charge" type and said central couplingshaft comprises an axial recess.
 9. The missile as claimed in claim 8,wherein said rear section of the missile comprises a compartment forstoring a parachute.
 10. The missile as claimed in claim 9, wherein thestabilizing tail unit is formed from a set of fins able to be foldedback against the body of the missile.
 11. A missile as claimed in claim5, forming a sub-projectile of a carrier projectile.
 12. A missile asclaimed in claim 6 forming a sub-projectile of a carrier projectile. 13.A missile as claimed in claim 7 forming a sub-projectile of a carrierprojectile.
 14. A missile as claimed in claim 8 forming a sub-projectileof a carrier projectile.
 15. A missile as claimed in claim 9 forming asub-projectile of a carrier projectile.
 16. A missile as claimed inclaim 10, forming a sub-projectile of a carrier projectile.